EP2343237B1 - Aircraft fuselage frame in composite material with stabilized web - Google Patents

Aircraft fuselage frame in composite material with stabilized web Download PDF

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Publication number
EP2343237B1
EP2343237B1 EP10382364.7A EP10382364A EP2343237B1 EP 2343237 B1 EP2343237 B1 EP 2343237B1 EP 10382364 A EP10382364 A EP 10382364A EP 2343237 B1 EP2343237 B1 EP 2343237B1
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EP
European Patent Office
Prior art keywords
frame
manufacture
segment
curing
manufacturing procedure
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Not-in-force
Application number
EP10382364.7A
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German (de)
French (fr)
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EP2343237A3 (en
EP2343237A2 (en
Inventor
Elena ARÉVALO RODRÍGUEZ
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Airbus Operations SL
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Airbus Operations SL
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Publication date
Priority to ES200931307A priority Critical patent/ES2383986B1/en
Application filed by Airbus Operations SL filed Critical Airbus Operations SL
Publication of EP2343237A2 publication Critical patent/EP2343237A2/en
Publication of EP2343237A3 publication Critical patent/EP2343237A3/en
Application granted granted Critical
Publication of EP2343237B1 publication Critical patent/EP2343237B1/en
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Classifications

    • BPERFORMING OPERATIONS; TRANSPORTING
    • B64AIRCRAFT; AVIATION; COSMONAUTICS
    • B64CAEROPLANES; HELICOPTERS
    • B64C1/00Fuselages; Constructional features common to fuselages, wings, stabilising surfaces and the like
    • B64C1/06Frames; Stringers; Longerons ; Fuselage sections
    • B64C1/061Frames
    • BPERFORMING OPERATIONS; TRANSPORTING
    • B29WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
    • B29DPRODUCING PARTICULAR ARTICLES FROM PLASTICS OR FROM SUBSTANCES IN A PLASTIC STATE
    • B29D99/00Subject matter not provided for in other groups of this subclass
    • B29D99/0003Producing profiled members, e.g. beams
    • BPERFORMING OPERATIONS; TRANSPORTING
    • B29WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
    • B29CSHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
    • B29C70/00Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
    • B29C70/04Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts comprising reinforcements only, e.g. self-reinforcing plastics
    • B29C70/28Shaping operations therefor
    • B29C70/40Shaping or impregnating by compression not applied
    • B29C70/42Shaping or impregnating by compression not applied for producing articles of definite length, i.e. discrete articles
    • B29C70/44Shaping or impregnating by compression not applied for producing articles of definite length, i.e. discrete articles using isostatic pressure, e.g. pressure difference-moulding, vacuum bag-moulding, autoclave-moulding or expanding rubber-moulding
    • B29C70/446Moulding structures having an axis of symmetry or at least one channel, e.g. tubular structures, frames
    • BPERFORMING OPERATIONS; TRANSPORTING
    • B29WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
    • B29CSHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
    • B29C70/00Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
    • B29C70/04Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts comprising reinforcements only, e.g. self-reinforcing plastics
    • B29C70/28Shaping operations therefor
    • B29C70/40Shaping or impregnating by compression not applied
    • B29C70/42Shaping or impregnating by compression not applied for producing articles of definite length, i.e. discrete articles
    • B29C70/46Shaping or impregnating by compression not applied for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs
    • BPERFORMING OPERATIONS; TRANSPORTING
    • B29WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
    • B29CSHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
    • B29C70/00Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
    • B29C70/04Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts comprising reinforcements only, e.g. self-reinforcing plastics
    • B29C70/28Shaping operations therefor
    • B29C70/40Shaping or impregnating by compression not applied
    • B29C70/42Shaping or impregnating by compression not applied for producing articles of definite length, i.e. discrete articles
    • B29C70/46Shaping or impregnating by compression not applied for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs
    • B29C70/48Shaping or impregnating by compression not applied for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs and impregnating the reinforcements in the closed mould, e.g. resin transfer moulding [RTM], e.g. by vacuum
    • BPERFORMING OPERATIONS; TRANSPORTING
    • B29WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
    • B29LINDEXING SCHEME ASSOCIATED WITH SUBCLASS B29C, RELATING TO PARTICULAR ARTICLES
    • B29L2031/00Other particular articles
    • B29L2031/30Vehicles, e.g. ships or aircraft, or body parts thereof
    • B29L2031/3076Aircrafts
    • B29L2031/3082Fuselages
    • BPERFORMING OPERATIONS; TRANSPORTING
    • B64AIRCRAFT; AVIATION; COSMONAUTICS
    • B64CAEROPLANES; HELICOPTERS
    • B64C1/00Fuselages; Constructional features common to fuselages, wings, stabilising surfaces and the like
    • B64C2001/0054Fuselage structures substantially made from particular materials
    • B64C2001/0072Fuselage structures substantially made from particular materials from composite materials
    • YGENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02TCLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00Aeronautics or air transport
    • Y02T50/40Weight reduction
    • YGENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10TTECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T428/00Stock material or miscellaneous articles
    • Y10T428/24Structurally defined web or sheet [e.g., overall dimension, etc.]
    • Y10T428/24174Structurally defined web or sheet [e.g., overall dimension, etc.] including sheet or component perpendicular to plane of web or sheet
    • YGENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10TTECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T428/00Stock material or miscellaneous articles
    • Y10T428/24Structurally defined web or sheet [e.g., overall dimension, etc.]
    • Y10T428/24174Structurally defined web or sheet [e.g., overall dimension, etc.] including sheet or component perpendicular to plane of web or sheet
    • Y10T428/24182Inward from edge of web or sheet
    • YGENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10TTECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T428/00Stock material or miscellaneous articles
    • Y10T428/24Structurally defined web or sheet [e.g., overall dimension, etc.]
    • Y10T428/24479Structurally defined web or sheet [e.g., overall dimension, etc.] including variation in thickness
    • Y10T428/24488Differential nonuniformity at margin
    • YGENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10TTECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T428/00Stock material or miscellaneous articles
    • Y10T428/24Structurally defined web or sheet [e.g., overall dimension, etc.]
    • Y10T428/24628Nonplanar uniform thickness material
    • YGENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10TTECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T428/00Stock material or miscellaneous articles
    • Y10T428/24Structurally defined web or sheet [e.g., overall dimension, etc.]
    • Y10T428/24628Nonplanar uniform thickness material
    • Y10T428/24661Forming, or cooperating to form cells

Description

    FIELD OF THE INVENTION
  • This invention refers to aircraft fuselage frames made of composite material, and more specifically to frames with stabilized webs.
  • BACKGROUND OF THE INVENTION
  • Load frames, in addition to shaping and stiffening the fuselage of an aircraft, are the structural elements in charge of withstanding and transferring the loads coming from other structural elements of the aircraft, e.g. the wings or the stabilizers.
  • Traditionally, load frames have been made of metallic material and with different sections, the most usual ones being C-, I- and J-sections which, through machining processes, form a ribbed framework that stabilizes the web of the frame.
  • In today's aeronautical industry, the strength-to-weight ratio is an element of utmost importance; therefore, there is a demand for frames manufactured or optimised with composite materials, primarily with carbon fibre, instead of metallic frames.
  • In the case of load frames, it is difficult to compete with machined metallic frames because, due to the high and differing stresses they must withstand, alternative frames in composite materials usually pose various manufacturing problems since their design is so complex. In any event, some proposals are now available in this respect, such as those described in the applications for patents WO 2008/092970 , US 2009/0026315 and WO 2009/030731 .
  • GB 2443542 A , which is the closest prior art document, discloses a fuselage structure that includes an upper fuselage portion defined by a first cross-section including at least three different radii of curvature and a lower fuselage portion defined by a second cross-section including at least three different radii of curvature, wherein the first cross-section is different from the second cross-section. The portions can be integral or joined together with a unitary floor. The bottom of the lower portion can be substantially flattened to provide additional rotational clearance for take-off and landing. The portions are of sandwich form with the bottom of the lower portion having an integral keel.
  • One of the aforementioned problems is related to the attempt to achieve a frame design in which the webs are stabilized, thus making it possible to optimise the weight. Although in the case of form frames there are now some proposals available, the same is not true of load frames in spite of the existing demand for them in the aeronautical industry.
  • This aim of this invention is to address this demand.
  • SUMMARY OF THE INVENTION
  • One object of this invention is to provide an aircraft fuselage frame with stabilized web made of composite material configured such that the strength-to-weight ratio is optimised.
  • Another object of this invention is to provide an aircraft fuselage frame with stabilized web made of composite material whose configuration facilitates the manufacture thereof.
  • First of all, these and other objects are achieved with a frame in which, in at least one of its segments, the cross section is a closed omega-shaped section formed by one head, two webs, two feet and a stiffening element between the two webs, and the web thickness is less in the span between the stiffening element and the head than in the span between the stiffening element and the feet. This facilitates the frame weight optimisation.
  • In one preferential embodiment, the frame has the aforementioned configuration along its entire length. This results in optimised frames that require stabilizing web along their entire length.
  • In another preferential embodiment, the frame can comprise several segments where one of them has the aforementioned configuration and the others a different configuration. This results in optimised frames for zones with different requirements.
  • Secondly, these and other objects are achieved by a frame manufacturing procedure whereby the manufacture of the aforementioned omega-shaped segment and an internal stiffening element comprises steps to:
    • Provide an external element whose cross section is a closed omega-shaped section formed by one head, two webs and two feet, and an internal element with closed omega-shaped section formed by one head, two webs and two feet, with their respective webs and feet oriented in parallel to those of the external element.
    • Join the internal element to the external element.
  • In one preferential embodiment, a reinforcement element is also provided to close off the corners between the internal element and the external element. This facilitates the manufacture of the frame by avoiding discontinuities at the corners between the internal element and the external element and the possible beginning of debonding, thus achieving a better structural performance of the whole.
  • In another preferential embodiment, the manufacturing procedure of the frame segment composed of an external element and an internal element comprises the following steps:
    • Manufacture and curing of the internal element preform, preferably with an RTM procedure (it is also feasible to do it with pre-impregnated material using a suitable tool).
    • Manufacture of the external element preform, preferably with a pre-impregnated material hot forming procedure.
    • Co-bonding of these preforms together in a curing cycle in autoclave.
  • In another preferential embodiment, the manufacturing procedure of the frame segment composed of an external element and an internal element comprises the following steps:
    • Manufacture and curing of the external element preform, preferably with an RTM procedure (it is also feasible to do it with pre-impregnated material using a suitable tool)
    • Manufacture of the internal element preform, preferably with a pre-impregnated material hot forming procedure.
    • Co-bonding of these preforms together in a curing cycle in autoclave.
  • In both cases, this provides a very efficient procedure for manufacturing the frame, since on one hand a great dimensional precision is achieved in the manufacture of the elements by RTM, which is a significant advantage particularly in the case of load frames in which the internal and external elements have different dimensions in different zones, and on the other hand it is an efficient procedure for manufacturing frame segments with high load stresses because it enhances the mechanical characteristics since the pre-impregnated material has better mechanical properties than the RTM material.
  • In another preferential embodiment, the manufacturing procedure of the frame segment composed of an external element and an internal element comprises the following steps:
    • Manufacture of the internal and external element preforms with a pre-impregnated material.
    • Co-curing of these preforms together in a curing cycle in autoclave.
  • This provides an efficient procedure for manufacturing frame segments with high load stresses and without significant dimensional variations.
  • In another preferential embodiment, the manufacturing procedure of the frame segment composed of an external element and an internal element comprises the following steps:
    • Manufacture of dry preforms of the internal and external elements.
    • Co-curing of these preforms together with an RTM process.
  • This provides an efficient procedure for manufacturing frame segments with average load stresses and with significant dimensional variations.
  • In another preferential embodiment, the manufacturing procedure of the frame segment composed of an external element and an internal element comprises the following steps:
    • Manufacture and curing of the internal and external elements separately.
    • Joining of these elements with adhesives.
  • This provides an efficient procedure for manufacturing frame segments subject to lower load requirements.
  • In another preferential embodiment, the manufacturing procedure of the frame segment composed of an external element and an internal element comprises the following steps:
    • Manufacture and curing of the internal and external elements separately.
    • Joining of these elements with rivets.
  • This provides an efficient procedure for manufacturing frame segments subject to lower load requirements.
  • Other characteristics and advantages of this invention will be deduced from the following detailed description of an embodiment illustrating the object of the invention in relation to the accompanying figures.
  • BRIEF DESCRIPTION OF THE FIGURES
    • Figure 1 shows the conventional stabilization solution of a frame web made of composite material with omega-shaped section.
    • Figure 2 shows a cross section of a frame as per this invention.
    • Figure 3 shows a schematic view of one embodiment of a frame as per this invention, structured in several segments of differing configuration.
    • Figures 4a, 4b and 4c show cross sections of different segments of the frame in Figure 3.
    • Figure 5a schematically shows the elements used to manufacture a frame according to the first embodiment of this invention, and Figure 5b shows the resulting frame.
    • Figure 6a schematically shows the elements used to manufacture a frame according to the second embodiment of this invention, and Figure 6b shows the resulting frame.
    • Figure 7 schematically shows one of the procedures to manufacture a frame as per this invention.
    DETAILED DESCRIPTION OF THE INVENTION
  • The application for patent WO 2008/092970 describes a load frame made of composite material with a closed section that can have an Π or omega shape made from three elements: two lateral elements and one base element, each made with an appropriate laminate to withstand the loads to which it is going to be subjected.
  • Whereas in a metallic frame with a similar configuration, it would be easy to machine the stiffening elements, the same is not true for a frame made of composite material. In a frame 3 of this kind, as shown in Figure 1, the conventional solution in composite material to stabilize its web and optimise the weight would be to join two L-shaped elements 5 to it.
  • The alternative as per this invention is a frame 9 in which, in at least one of its segments, the cross section 11 (see Figure 2) is a closed omega-shaped section formed by one head 23, two webs 25, 25', two feet 27, 27' and a stiffening element 29 between the two webs 25, 25'. This cross section facilitates optimisation of the frame weight.
  • The head 23, the webs 25, 25', the feet 27, 27' and the stiffening element 29 are formed by layers of composite material arranged at 0°, +/-45° and 90°. The composite material may either be carbon fibre or glass fibre with thermo-setting or thermo-plastic resin.
  • Also included in the feet 27, 27', the head 23 and the stiffening element 29 are unidirectional fibre reinforcements at 0°, made of the same material or a compatible material and extending longitudinally along the entire frame without discontinuity. The reinforcement material has a high elastic module such that it makes the feet 27, 27', the head 23 and the stiffening element 29 having a high stiffness.
  • The webs 25, 25' can also have unidirectional fibre reinforcements of the same material or a compatible material, in any direction in which they can be continuous along the entire frame or local, depending on the stresses to which it is subjected. This means that the webs 25, 25' are capable of withstanding much greater loads than if they were composed exclusively of materials at 0°, +/-45° and 90°.
  • In one preferential embodiment of this invention, all of the frames 9 joined to the coating 17 of the fuselage, provides the cross section 11 shown in Figure 2.
  • In another preferential embodiment of this invention, illustrated in Figures 3 and 4, the frame 9 joined to the coating 17 of the fuselage can have, in addition to a cross section 11 in the first segment 31 (according to B-B') in a stiffened omega shape (illustrated in Figures 2 and 4c), different cross sections in other segments, e.g. a cross section 13 (according to A-A') in Π shape (illustrated in Figure 4a) in the second segment 33, or a cross section 15 (according to C-C') in a simple omega shape (illustrated in Figure 4b) in the third segment 35.
  • The second segment 33 corresponds to a segment of the frame 9 with local load inputs via fittings that are joined to the frame webs. Therefore, in this segment, a section with a stiffening element is not required and the frame 9 can be formed solely by an element with closed Π-shaped section (it could also be a simple omega shape) because, since the fittings are located in this zone, the webs do not have to be stabilized.
  • The third segment 35 corresponds to a segment of the frame with lower stresses and, therefore, a simple omega-shaped cross section 15 would be appropriate.
  • Between the aforementioned segments, there are logically transition zones between the differently shaped sections.
  • In any of its embodiments, the frame according to this invention is applicable to circular, ellipsoidal, rectangular and other shapes of fuselage sections.
  • According to a first embodiment of the manufacturing procedure under this invention, and as illustrated in Figures 5a and 5b, the first segment 31 of the frame with the cross section 11 is manufactured by joining an internal element 51 with closed omega-shaped section, formed by one head 53, two webs 55, 55' and two feet 57, 57', to an external element 41 with closed omega-shaped section, formed by one head 43, two webs 45, 45' and two feet 47, 47'.
  • Thus the webs 25, 25' of the resulting frame have a span with the thickness resulting from joining the webs 55, 55' of the internal element 51 and the webs 45, 45' of the external element 41, and another span with the thickness of the webs 45, 45' of the external element 41. The feet 27, 27 of the resulting frame have the thickness resulting from joining the feet 57, 57' of the internal element 51 and the feet 47, 47' of the external element 41. Finally, the stiffening element 29 corresponds to the head 53 of the internal element 51. This cross section facilitates optimisation of the frame weight and provides a necessary reinforcement in the zone that is left between stringers mouseholes ("castellation").
  • According to a second embodiment of the manufacturing procedure under this invention, and as illustrated in Figures 6a and 6b, the first segment 31 of the frame with the cross section 11 is manufactured using a third element 61 formed with layers of composite material to close off the corners between the internal element 51 and the external element 41. The central zone 61 of this element is joined to the head 53 of the internal element 51 and the end zones 65, 65' to the webs 45, 45' of the external element 41.
  • Following is a description of a first embodiment variant of the procedure, according to the invention, to manufacture a segment of a frame with the internal element 51 and the external element 41.
  • In a first stage 71, a preform 51' of the internal element 51 would be manufactured and would be cured using an RTM (Resin Transfer Moulding) process (it is also feasible to do it with pre-impregnated material using a suitable tool). It is well known that this process uses a closed, pressurized mould in which dry preforms are placed and then resin is injected.
  • In a second stage, a preform 41' of the external element 41 would be manufactured using a hot forming process, with a first step 75 of pre-impregnated material (prepeg) stacking and a second step 77 of hot forming.
  • In a third stage 79, the preforms 51' and 41' would be co-bonded together and the piece would be consolidated in a curing cycle in autoclave. As seen in Figure 7, this process requires, on one hand, a first external tool 81 on which the preform 41' would lay and, on the other hand, an internal tool 83 and a second external tool 85 to guarantee adequate compacting of the preform 41'.
  • By manufacturing the external element 41 and the internal element 51 separately, each with variations of thickness and section, the frame obtained from joining them is optimised, thus achieving a variable closed section.
  • In a second embodiment variant of a procedure, according to the invention, to manufacture a segment of a frame with the aforementioned internal element 51 and external element 41, a preform 41' of the external element 41 would be manufactured and would be cured using an RTM process (it is also feasible to do it with pre-impregnated material using a suitable tool). On the other hand, a preform 51' of the internal element 51 would be manufactured using a hot forming process, and finally the preforms 51' and 41' would be co-bonded together and the piece would be consolidated in a curing cycle in autoclave.
  • In a third embodiment variant of a procedure, according to the invention, to manufacture a segment of a frame with the aforementioned internal element 51 and external element 41, their preforms 51', 41' would be manufactured separately and, after a process of hot forming and duly arranged in appropriate tools, they would be co-cured together in a curing cycle in autoclave.
  • In a fourth embodiment variant of a procedure, according to the invention, to manufacture a segment of a frame with the aforementioned internal element 51 and external element 41, their dry preforms 51', 41' would be manufactured separately and they would be co-cured together by an RTM process.
  • In a fifth embodiment variant of a procedure, according to the invention, to manufacture a segment of a frame with the aforementioned internal element 51 and external element 41, these elements would be manufactured and cured separately and they would be joined with adhesives.
  • In a sixth embodiment variant of a procedure, according to the invention, to manufacture a segment of a frame with the aforementioned internal element 51 and external element 41, these elements would be manufactured and cured separately and they would be joined with rivets.
  • Although this invention has been completely described in connection with the preferred embodiments, it is obvious that any modifications within its scope can be introduced, as the scope is not considered to be restricted by the preceding embodiments, but rather by the contents of the following claims.

Claims (18)

  1. Frame (9) for the fuselage of an aircraft made of composite material, in which at least in a first segment (31) of the frame its cross section (11) is a closed omega-shaped section formed by one head (23), two webs (25, 25') and two feet (27, 27') with a stiffening element (29) between the two webs (25, 25'), characterized in that the thickness of the webs (25, 25') is less in the span between the stiffening element (29) and the head (23) than in the span between the stiffening element (29) and the feet (27, 27').
  2. Frame (9) according to claim 1, characterized in that this first segment (31) extends along its entire length.
  3. Frame (9) according to claim 1, characterized in that, in addition to the first segment (31), it also comprises a second segment (33) in the zone provided for reception of external loads, whose cross section is a closed section.
  4. Frame (9) according to claim 3, characterized in that this second segment (33) has an Π-shaped cross section (13).
  5. Frame (9) according to claim 3, characterized in that this second segment (33) has an omega-shaped cross section (15).
  6. Frame (9) according to claim 1, characterized in that, in addition to the first segment (31), it also comprises a third segment (35) that has an omega-shaped cross section (15).
  7. Manufacturing procedure for a frame (9) according to any of claims 1-6, characterized in that the manufacture of the first segment (31) comprises steps to:
    a) provide an external element (41) whose cross section is a closed omega-shaped section formed by one head (43), two webs (45, 45') and two feet (47, 47'), and an internal element (51) with closed omega-shaped section formed by one head (53), two webs (55, 55') and two feet (57, 57'), with their respective webs (45, 55; 45', 45') and feet (47, 57; 47', 57') oriented in parallel;
    b) join the internal element (51) to the external element (41).
  8. Manufacturing procedure for a frame (9) according to claim 7, characterized in that a reinforcement element (61) is also provided to close off the corners between the internal element (51) and the external element (41).
  9. Manufacturing procedure for a frame (9) according to claim 7, characterized in that the manufacture of the first segment (31) comprises the following steps:
    a) manufacture and curing of the preform (51') of the internal element (51);
    b) manufacture of the preform (41') of the external element (41);
    c) co-bonding of these preforms (41', 51') together in a curing cycle in autoclave.
  10. Manufacturing procedure for a frame (9) according to claim 9, characterized in that stage a) is carried out with a resin transfer moulding (RTM) procedure.
  11. Manufacturing procedure for a frame (9) according to any of claims 9-10, characterized in that stage b) is carried out with pre-impregnated material by a hot forming procedure.
  12. Manufacturing procedure for a frame (9) according to claim 7 characterized in that the manufacture of the first segment (31) comprises the following steps:
    a) manufacture of the preform (51') of the internal element (51);
    b) manufacture and curing of the preform (41') of the external element (41);
    c) co-bonding of these preforms (41', 51') together in a curing cycle in autoclave.
  13. Manufacturing procedure for a frame (9) according to claim 12, characterized in that stage a) is carried out with pre-impregnated material by a hot forming procedure.
  14. Manufacturing procedure for a frame (9) according to any of claims 12-13, characterized in that stage b) is carried out with a resin transfer moulding (RTM) procedure.
  15. Manufacturing procedure for a frame (9) according to claim 7, characterized in that the manufacture of the first segment (31) comprises the following steps:
    a) manufacture of a preform (51') of the internal element (51) in pre-impregnated material;
    b) manufacture of a preform (41') of the external element (41) in pre-impregnated material;
    c) co-curing of these preforms (41', 51') together in a curing cycle in autoclave.
  16. Manufacturing procedure for a frame (9) according to claim 7, characterized in that the manufacture of the first segment (31) comprises the following steps:
    a) manufacture of a dry preform (51') of the internal element (51);
    b) manufacture of a dry preform (41') of the external element (41);
    c) co-curing of these preforms (41', 51') together with a resin transfer moulding (RTM) process.
  17. Manufacturing procedure for a frame (9) according to claim 7, characterized in that the manufacture of the first segment (31) comprises the following steps:
    a) manufacture and curing of the internal element (51);
    b) manufacture and curing of the external element (41);
    c) joining of the internal element (51) and the external element (41) with adhesives.
  18. Manufacturing procedure for a frame (9) according to claim 7, characterized in that the manufacture of the first segment (31) comprises the following steps:
    a) manufacture and curing of the internal element (51);
    b) manufacture and curing of the external element (41);
    c) joining of the internal element (51) and the external element (41) with rivets.
EP10382364.7A 2009-12-30 2010-12-30 Aircraft fuselage frame in composite material with stabilized web Not-in-force EP2343237B1 (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
ES200931307A ES2383986B1 (en) 2009-12-30 2009-12-30 AIRCRAFT FUSELAGE NOTEBOOK IN COMPOSITE MATERIAL WITH STABILIZED SOUL.

Publications (3)

Publication Number Publication Date
EP2343237A2 EP2343237A2 (en) 2011-07-13
EP2343237A3 EP2343237A3 (en) 2013-07-24
EP2343237B1 true EP2343237B1 (en) 2014-12-17

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EP10382364.7A Not-in-force EP2343237B1 (en) 2009-12-30 2010-12-30 Aircraft fuselage frame in composite material with stabilized web

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US (1) US8597771B2 (en)
EP (1) EP2343237B1 (en)
BR (1) BRPI1010451A2 (en)
CA (1) CA2726594C (en)
ES (1) ES2383986B1 (en)

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US10919260B2 (en) * 2019-05-09 2021-02-16 The Boeing Company Composite structure having a variable gage and methods for forming a composite structure having a variable gage
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Also Published As

Publication number Publication date
US8597771B2 (en) 2013-12-03
ES2383986A1 (en) 2012-06-28
ES2383986B1 (en) 2013-05-16
CA2726594A1 (en) 2011-06-30
BRPI1010451A2 (en) 2013-03-19
US20110159242A1 (en) 2011-06-30
EP2343237A3 (en) 2013-07-24
CA2726594C (en) 2017-03-14
EP2343237A2 (en) 2011-07-13

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